This invention relates generally to gas turbine engines and more particularly to stationary flowpath components, such as nozzle segments and turbine shrouds, used in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. Each turbine stage commonly includes a turbine rotor and a stationary turbine nozzle for channeling combustion gases into the turbine rotor disposed downstream thereof. The turbine rotor includes a plurality of circumferentially spaced apart blades extending radially outwardly from a rotor disk that rotates about the centerline axis of the engine. The nozzle includes a plurality of circumferentially spaced apart vanes radially aligned with the rotor blades. Turbine nozzles are typically segmented around the circumference thereof with each nozzle segment having one or more nozzle vanes disposed between inner and outer bands that define the radial flowpath boundaries for the hot combustion gases flowing through the nozzle.
Each turbine stage further includes a shroud assembly located immediately downstream of the turbine nozzle. The shroud assembly closely surrounds the turbine rotor and thus defines the outer radial flowpath boundary for the hot combustion gases flowing therethrough. A typical shroud assembly comprises a shroud support which is fastened to the engine outer case and which in turn supports a plurality of shrouds. The shrouds are arcuate, rectangular components arranged circumferentially in an annular array so as to encircle the rotor blades.
A small gap exists between the adjoining circumferential edges of adjacent shrouds. To reduce leakage between adjacent shrouds, the gaps are sealed with spline seals, which comprise thin sheet metal strips inserted into slots cut in the circumferential edges of the shrouds so as to span the gap. Similarly, both the outer and inner bands of each nozzle segment define circumferential edges that adjoin the circumferential edges of adjacent nozzle segments. These junctions are likewise sealed with spline seals.
To better accommodate the spline seals, the circumferential edges of the shrouds and nozzle segments have been limited to linear edges. Typically, nozzle circumferential edges form either a single straight cut at an angle relative to the axial direction of the engine, or a series of straight cuts arranged in a "dog leg" or "z-shaped" configuration to reduce overhang with respect to the airfoil. Shrouds are generally rectangular with their circumferential edges running in the axial direction.
Because nozzle segments and shrouds operate in a high temperature environment, it is necessary that they be cooled to avoid reduced service life or even material failure. This cooling is ordinarily accomplished internally by using cooling air extracted from the compressor. However, the linear circumferential edges of these stationary flowpath components result in sharp corners that are difficult to cool. This is particularly so with the single cut angled edges of many nozzle segments, which define an extremely sharp corner. Moreover, the linear circumferential edges of the nozzle segments and the shrouds do not follow the naturally curved flow lines of the nozzle and the turbine rotor, respectively. Thus, the gases cross each gap between adjoining circumferential edges (in both the nozzle and the rotor) at least once. When this occurs, the boundary layer of the gases is "tripped," thereby causing turbulence and higher convection coefficients, which results in local areas of higher temperature along the circumferential edges. As a result of boundary layer tripping, the circumferential edges of nozzle segments and shrouds are often the most difficult areas to cool.
Accordingly, there is a need for stationary flowpath components that avoid boundary layer tripping.